Adiabatic Film Cooling Effectiveness of a LAM Fabricated Porous Leading Edge Segment of a Turbine Blade

2018 ◽  
Author(s):  
Luisana Calderon ◽  
Andres Curbelo ◽  
Gaurav Gupta ◽  
Jayanta S. Kapat
Keyword(s):  
Additive Manufacturing ◽  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Metal Powders ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Porous Region ◽  
External Cooling

Laser additive manufacturing (LAM) is an emerging technology that builds parts in a layer-by-layer process by selectively melting metal powders. This additive manufacturing technique among others can produce very complex geometries, which are not possible using conventional methods. A mock segment of the leading edge of a turbine blade, designed with both internal and external cooling features fabricated by LAM of Inconel powder, is investigated. This design consists of an internal impingement cooling array and an engineered-porous structure. This porous region consists of a lattice of intersecting cylinders that simulates the effect of a transpiration cooled segment or permeable wall with a designed porosity of 0.57. Transpiration cooling is a promising external cooling technique capable of reducing thermal gradients at the surface of the blade by providing a more uniform film than conventional discrete film cooling holes. In this current study, adiabatic film cooling effectiveness is experimentally investigated using pressure sensitive paint (PSP) for blowing ratios ranging between 0.03 and 0.15. Using air as the mainstream, and CO2 as the coolant source, a density ratio of 1.5 is obtained. Steady state simulations using RANS are analyzed and used to compare against experimental results. All cases result in an increase in effectiveness values with increasing blowing ratio. Highest effectiveness values are associated with high pressure drop. Coolant uniformity is observed downstream the porous region and mixing becomes more significant as blowing ratio is increased.

Numerical Analysis on the Leading Edge Film Cooling of Bifurcation Holes for Gas Turbine Blade

2021 ◽  
Author(s):  
Zhonghao Tang ◽  
Gongnan Xie ◽  
Honglin Li ◽  
Wenjing Gao ◽  
Chunlong Tan ◽  
...  
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Turbulence Models ◽  
Leading Edge ◽  
Spanwise Direction ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Cylindrical Holes ◽  
Lift Off

Abstract Film cooling performance of the cylindrical film holes and the bifurcated film holes on the leading edge model of the turbine blade are investigated in this paper. The suitability of different turbulence models to predict local and average film cooling effectiveness is validated by comparing with available experimental results. Three rows of holes are arranged in a semi-cylindrical model to simulate the leading edge of the turbine blade. Four different film cooling structures (including a cylindrical film holes and other three different bifurcated film holes) and four different blowing ratios are studied in detail. The results show that the film jets lift off gradually in the leading edge area as the blowing ratio increases. And the trajectory of the film jets gradually deviate from the mainstream direction to the spanwise direction. The cylindrical film holes and vertical bifurcated film holes have better film cooling effectiveness at low blowing ratio while the other two transverse bifurcated film holes have better film cooling effectiveness at high blowing ratio. And the film cooling effectiveness of the transverse bifurcated film holes increase with the increasing the blowing ratio. Additionally, the advantage of transverse bifurcated holes in film cooling effectiveness is more obvious in the downstream region relative to the cylindrical holes. The Area-Average film cooling effectiveness of transverse bifurcated film holes is 38% higher than that of cylindrical holes when blowing ratio is 2.

Enhancement of Film Cooling Performance at the Leading Edge of Turbine Blade

2006 ◽  
Author(s):  
K.-S. Kim ◽  
Youn J. Kim ◽  
S.-M. Kim
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Cylindrical Body ◽  
Spanwise Direction ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio

To enhance the film cooling performance in the vicinity of the turbine blade leading edge, the flow characteristics of the film-cooled turbine blade have been investigated using a cylindrical body model. The inclination of the cooling holes is along the radius of the cylindrical wall and 20 deg relative to the spanwise direction. Mainstream Reynolds number based on the cylinder diameter was 1.01×105 and 0.69×105, and the mainstream turbulence intensities were about 0.2% in both Reynolds numbers. CO2 was used as coolant to simulate the effect of density ratio of coolant-to-mainstream. Furthermore, the effect of coolant flow rates was studied for various blowing ratios of 0.4, 0.7, 1.1, and 1.4, respectively. In experiment, spatially-resolved temperature distributions along the cylindrical body surface were visualized using infrared thermography (IRT) in conjunction with thermocouples, digital image processing, and in situ calibration procedures. This comparison shows the results generated to be reasonable and physically meaningful. The film cooling effectiveness of current measurement (0.29 mm × 0.33 min per pixel) presents high spatial and temperature resolutions compared to other studies. Results show that the blowing ratio has a strong effect on film cooling effectiveness and the coolant trajectory is sensitive to the blowing ratio. The local spanwise-averaged effectiveness can be improved by locating the first-row holes near the second-row holes.

Simulation Film Cooling in the Leading Edge Region of a Turbine Blade (Trench Effect on Film Effectiveness from Cylinder in Crossflow)

2014 ◽  
Vol 554 ◽  
pp. 317-321
Author(s):  
Mohamad Rasidi Bin Pairan ◽  
Norzelawati Binti Asmuin ◽  
Hamidon bin Salleh
Keyword(s):  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Edge Region ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Hot Gas ◽  
Cooling Technique

Film cooling is one of the cooling techniques applied to the turbine blade. Gas turbine used film cooling technique to protect turbine blade from directly expose to the hot gas to avoid the blade from defect. The focus of this investigation is to investigate the effect of embedded three difference depth of trench at coolant holes geometry. Comparisons are made at four difference blowing ratios which are 1.0, 1.25 and 1.5. Three configuration leading edge with depth Case A (0.0125D), Case B (0.0350D) and Case C (0.713D) were compared to leading edge without trench. Result shows that as blowing ratio increased from 1.0 to 1.25, the film cooling effectiveness is increase for leading edge without trench and also for all cases. However when the blowing ratio is increase to 1.5, film cooling effectiveness is decrease for all cases. Overall the Case B with blowing ratio 1.25 has the best film cooling effectiveness with significant improvement compared to leading edge without trench and with trench Case A and Case C.

Prediction of Film Cooling Effectiveness on a Gas Turbine Blade Leading Edge Using ANN and CFD

2018 ◽  
Vol 35 (2) ◽  
pp. 101-111 ◽  
Author(s):  
J. O. Dávalos ◽  
J. C. García ◽  
G. Urquiza ◽  
A. Huicochea ◽  
O. De Santiago
Keyword(s):  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Hole Diameter ◽  
Relative Importance ◽  
Gas Turbine Blade ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio

Abstract In this work, the area-averaged film cooling effectiveness (AAFCE) on a gas turbine blade leading edge was predicted by employing an artificial neural network (ANN) using as input variables: hole diameter, injection angle, blowing ratio, hole and columns pitch. The database used to train the network was built using computational fluid dynamics (CFD) based on a two level full factorial design of experiments. The CFD numerical model was validated with an experimental rig, where a first stage blade of a gas turbine was represented by a cylindrical specimen. The ANN architecture was composed of three layers with four neurons in hidden layer and Levenberg-Marquardt was selected as ANN optimization algorithm. The AAFCE was successfully predicted by the ANN with a regression coefficient R2<0.99 and a root mean square error RMSE=0.0038. The ANN weight coefficients were used to estimate the relative importance of the input parameters. Blowing ratio was the most influential parameter with relative importance of 40.36 % followed by hole diameter. Additionally, by using the ANN model, the relationship between input parameters was analyzed.

Heat Transfer Coefficients and Film Cooling Effectiveness on the Squealer Tip of a Gas Turbine Blade

2003 ◽  
Vol 125 (4) ◽  
pp. 648-657 ◽  
Author(s):  
Jae Su Kwak ◽  
Je-Chin Han
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Rotor Blade ◽  
Transfer Coefficients ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Heat Transfer Coefficients ◽  
Blowing Ratio

Experimental investigations were performed to measure the detailed heat transfer coefficients and film cooling effectiveness on the squealer tip of a gas turbine blade in a five-bladed linear cascade. The blade was a two-dimensional model of a first stage gas turbine rotor blade with a profile of the GE-E3 aircraft gas turbine engine rotor blade. The test blade had a squealer (recessed) tip with a 4.22% recess. The blade model was equipped with a single row of film cooling holes on the pressure side near the tip region and the tip surface along the camber line. Hue detection based transient liquid crystals technique was used to measure heat transfer coefficients and film cooling effectiveness. All measurements were done for the three tip gap clearances of 1.0%, 1.5%, and 2.5% of blade span at the two blowing ratios of 1.0 and 2.0. The Reynolds number based on cascade exit velocity and axial chord length was 1.1×106 and the total turning angle of the blade was 97.9 deg. The overall pressure ratio was 1.2 and the inlet and exit Mach numbers were 0.25 and 0.59, respectively. The turbulence intensity level at the cascade inlet was 9.7%. Results showed that the overall heat transfer coefficients increased with increasing tip gap clearance, but decreased with increasing blowing ratio. However, the overall film cooling effectiveness increased with increasing blowing ratio. Results also showed that the overall film cooling effectiveness increased but heat transfer coefficients decreased for the squealer tip when compared to the plane tip at the same tip gap clearance and blowing ratio conditions.

Influence of Turbine Blade Leading Edge Profile on Film Cooling With Shaped Holes

2017 ◽  
Author(s):  
Mingjie Zhang ◽  
Nian Wang ◽  
Andrew F. Chen ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Stagnation Line ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Edge Profile ◽  
Elliptical Cylinders ◽  
Blade Leading Edge

This paper presents the turbine blade leading edge model film cooling effectiveness with shaped holes, using the pressure sensitive paint (PSP) mass transfer analogy method. The effects of leading edge profile, coolant to mainstream density ratio and blowing ratio are studied. Computational simulations are performed using the realizable k-ε turbulence model. Effectiveness obtained by CFD simulations are compared with experiments. Three leading edge profiles, including one semi-cylinder and two semi-elliptical cylinders with an after body, are investigated. The ratios of major to minor axis of two semi-elliptical cylinders are 1.5 and 2.0, respectively. The leading edge has three rows of shaped holes. For the semi-cylinder model, shaped holes are located at 0 degrees (stagnation line) and ± 30 degrees. Row spacing between cooling holes and the distance between impingement plate and stagnation line are the same for three leading edge models. The coolant to mainstream density ratio varies from 1.0 to 1.5 and 2.0, and the blowing ratio varies from 0.5 to 1.0 and 1.5. Mainstream Reynolds number is about 100,900 based on the diameter of the leading edge cylinder, and the mainstream turbulence intensity is about 7%. The results provide an understanding of the effects of leading edge profile and on turbine blade leading edge region film cooling with shaped-hole designs.

Numerical Study on the Influence of Trench Width on Film Cooling Characteristics of Double-Wave Trench

2017 ◽  
Author(s):  
Bo-lun Zhang ◽  
Li Zhang ◽  
Hui-ren Zhu ◽  
Jian-sheng Wei ◽  
Zhong-yi Fu
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Numerical Study ◽  
Leading Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Trench Width

Film cooling performance of the double-wave trench was numerically studied to improve the film cooling characteristics. Double-wave trench was formed by changing the leading edge and trailing edge of transverse trench into cosine wave. The film cooling characteristics of transverse trench and double-wave trench were numerically studied using Reynolds Averaged Navier Stokes (RANS) simulations with realizable k-ε turbulence model and enhanced wall treatment. The film cooling effectiveness and heat transfer coefficient of double-wave trench at different trench width (W = 0.8D, 1.4D, 2.1D) conditions are investigated, and the distribution of temperature field and flow field were analyzed. The results show that double-wave trench effectively improves the film cooling effectiveness and the uniformity of jet at the downstream wall of the trench. The span-wise averaged film cooling effectiveness of the double-wave trench model increases 20–63% comparing with that of the transverse trench at high blowing ratio. The anti-counter-rotating vortices which can press the film on near-wall are formed at the downstream wall of the double-wave trench. With the double-wave trench width decreasing, the film cooling effectiveness gradually reduces at the hole center-line region of the downstream trench. With the increase of the blowing ratio, the span-wise averaged heat transfer coefficient increases. The span-wise averaged heat transfer coefficient of the double-wave trench with 0.8D and 2.1D trench width is higher than that of the double-wave trench with 1.4D trench width at the high blowing ratio conditions.

Leading Edge Film Cooling: Computations and Experiments Including Density Effects

1996 ◽  
Author(s):  
Pingfan He ◽  
Dragos Licu ◽  
Martha Salcudean ◽  
Ian S. Gartshore
Keyword(s):  
Experimental Data ◽  
Film Cooling ◽  
Stokes Equations ◽  
Leading Edge ◽  
Variable Density ◽  
Navier Stokes ◽  
Cooling Effectiveness ◽  
Stagnation Line ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio

The effect of varying coolant density on film cooling effectiveness for a turbine blade-model was numerically investigated and compared with experimental data. This model had a semi-circular leading edge with four rows of laterally-inclined film cooling orifices positioned symmetrically about the stagnation line. A curvilinear coordinate-based CFD code was developed and used for the numerical investigation. The code used a domain segmentation strategy in conjunction with general curvilinear grids to model the complex blade configuration. A multigrid method was used to accelerate the convergence rate. The time-averaged, variable-density, Navier-Stokes equations together with the energy or scalar equation were solved. Turbulence closure was attained by the standard k–ε model with a near-wall k model. Either air or CO2 was used as coolant in three cases of injection through single rows and alternatively staggered double raws of holes. Two different blowing rates were investigated in each case and compared with experimental data. The experimental results were obtained using a wind tunnel model, and the mass/heat analogy was used to determine the film cooling effectiveness. The higher density of the carbon dioxide coolant (approximately 1.5 times the density of air) in the isothermal mass injection experiments, was used to simulate the effects of injection of a colder air in the corresponding adiabatic heat transfer situation. Good agreement between calculated and measured film cooling effectiveness was found for low blowing ratio M ≤ 0.5 and the effect of density was not significant. At higher blowing ratio M > 1 the calculations consistently overpredict the measured values of film cooling effectiveness.

Turbine Blade Platform Film Cooling With Simulated Swirl Purge Flow and Slashface Leakage Conditions

2016 ◽  
Author(s):  
Andrew F. Chen ◽  
Chao-Cheng Shiau ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Suction Side ◽  
Leakage Flow ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Purge Flow ◽  
Endwall Film Cooling ◽  
Film Cooling Holes

The combined effects of inlet purge flow and the slashface leakage flow on the film cooling effectiveness of a turbine blade platform were studied using the pressure sensitive paint (PSP) technique. Detailed film cooling effectiveness distributions on the endwall were obtained and analyzed. The inlet purge flow was generated by a row of equally-spaced cylindrical injection holes inside a single-tooth generic stator-rotor seal. In addition to the traditional 90 degree (radial outward) injection for the inlet purge flow, injection at a 45 degree angle was adopted to create a circumferential/azimuthal velocity component toward the suction side of the blades, which created a swirl ratio (SR) of 0.6. Discrete cylindrical film cooling holes were arranged to achieve an improved coverage on the endwall. Backward injection was attempted by placing backward injection holes near the pressure side leading edge portion. Slashface leakage flow was simulated by equally-spaced cylindrical injection holes inside a slot. Experiments were done in a five-blade linear cascade with an average turbulence intensity of 10.5%. The inlet and exit Mach numbers were 0.26 and 0.43, respectively. The inlet and exit mainstream Reynolds numbers based on the axial chord length of the blade were 475,000 and 720,000, respectively. The coolant-to-mainstream mass flow ratios (MFR) were varied from 0.5%, 0.75%, to 1% for the inlet purge flow. For the endwall film cooling holes and slashface leakage flow, blowing ratios (M) of 0.5, 1.0, and 1.5 were examined. Coolant-to-mainstream density ratios (DR) that range from 1.0 (close to low temperature experiments) to 1.5 (intermediate DR) and 2.0 (close to engine conditions) were also examined. The results provide the gas turbine engine designers a better insight into improved film cooling hole configurations as well as various parametric effects on endwall film cooling when the inlet (swirl) purge flow and slashface leakage flow were incorporated.

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